U.S. patent number 5,963,182 [Application Number 08/888,762] was granted by the patent office on 1999-10-05 for edge-supported umbrella reflector with low stowage profile.
Invention is credited to Samir F. Bassily.
United States Patent |
5,963,182 |
Bassily |
October 5, 1999 |
Edge-supported umbrella reflector with low stowage profile
Abstract
An umbrella-like antenna reflector assembly for use on an
orbiting spacecraft. The reflector has a main rib and a plurality
of secondary ribs each connected to a hub assembly by a respective
hinge mechanism such that activation of the hub assembly causes the
reflector to move between collapsed and opened configurations. The
reflector further has a mesh member attached to the ribs. A
deployment boom connects the main rib of the reflector to the
spacecraft. The deployment boom is operable with the main rib and
the spacecraft to move the reflector between a collapsed and stowed
configuration proximate the spacecraft and an open and deployed
configuration outside the spacecraft. The storage profile is
sufficiently slim to permit launching of a 6-25 meter diameter
reflector attached to a full-sized spacecraft on one or more
commercially available launch vehicles without the need for mid-rib
hinges. A feed assembly is connected to the spacecraft. The feed
assembly is offset from and operable with the mesh member of the
reflector when the reflector is in the opened and deployed
configurations to receive and/or transmit radio frequency energy
therefrom.
Inventors: |
Bassily; Samir F. (Los Angeles,
CA) |
Family
ID: |
25393836 |
Appl.
No.: |
08/888,762 |
Filed: |
July 7, 1997 |
Current U.S.
Class: |
343/912; 343/881;
343/882; 343/915 |
Current CPC
Class: |
H01Q
15/161 (20130101); H01Q 1/288 (20130101) |
Current International
Class: |
H01Q
15/14 (20060101); H01Q 1/28 (20060101); H01Q
15/16 (20060101); H01Q 1/27 (20060101); H01Q
015/20 () |
Field of
Search: |
;343/915,912,840,916,913,914,721,894 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Kim; Robert H.
Assistant Examiner: Lauchman; Layla
Attorney, Agent or Firm: Gudmestad; Terje Grunebach;
Georgann Sales; Michael W.
Parent Case Text
CROSS-REFERENCE TO RELATED APPLICATIONS
The present application is related to U.S. patent application Ser.
No. 08/888,487, entitled "A Continually Adjustable Nonreturn Knot,"
attorney docket No. PD-970125, U.S. patent application Ser. No.
08/888,486, entitled "Mesh Tensioning, Retension And Management
Systems For Large Deployable Reflectors," attorney docket No.
PD-960515, U.S. patent application Ser. No. 08/888,485, entitled
"High-Torque Apparatus and Method Using Composite Materials For
Deployment Of A Multi-Rib Umbrella-Type Reflector," attorney docket
No. PD-970182, and U.S. patent application Ser. No. 08/888,500,
entitled "Apparatus And Method For Combined Redundant Deployment
And Launch Locking Of Deployable Satellite Appendages," attorney
docket No. PD-960507. The present application and the four related
applications, which are incorporated herein by reference, were
filed with the U.S. Patent Office on the same day.
Claims
What is claimed is:
1. For use on an orbiting spacecraft, a reflector antenna system
comprising:
an umbrella-like reflector having a main rib and a plurality of
secondary ribs each connected to a hub assembly by a respective
hinge mechanism such that activation of the hub assembly causes the
reflector to move between collapsed and opened configurations, the
reflector further having a mesh member attached to said main rib
and said plurality of secondary ribs;
a deployment boom connecting the main rib of the reflector of the
spacecraft, wherein the deployment boom is operable with the main
rib and the spacecraft to move the reflector between a collapsed
and stowed configuration adjacent to the spacecraft and a deployed
configuration away from the spacecraft, wherein the total number of
ribs is an odd number such that the deployment boom can be
positioned at least partially between a pair of secondary ribs
situated opposite from the main rib when the reflector is in the
collapsed and stowed configuration; and
a feed assembly connected to the spacecraft, the feed assembly
being offset from and operable with said mesh member of the
reflector when the reflector is in the deployed configuration to
transmit radio frequency energy therefrom.
2. The reflector antenna system of claim 1 further comprising two
opposing hinge straps connecting each of said ribs to said hub
assembly.
3. The reflector antenna system of claim 1 wherein said deployment
boom is kinked to afford low profile stowage of the reflector in
the collapsed and stowed configurations.
4. The reflector antenna system of claim 1 wherein the main rib
consists of an inner main rib and an outer main rib spliced to said
inner main rib.
5. The reflector antenna system of claim 1 further comprising a
network of pretensioned radial and circumferential retention chords
associated with said mesh member to resist the natural pillowing
tendency of said mesh member.
6. The reflector antenna system of claim 5 wherein the
circumferential spacing of the ribs varies from rib to rib to
minimize mesh faceting errors.
7. The reflector antenna system of claim 5 wherein said mesh member
is attached to the ribs at radial attachment points, wherein the
spacing of the radial attachment points decreases as the
circumference of the reflector increase to minimize mesh faceting
errors.
8. The reflector antenna system of claim 1 wherein said hub
assembly includes a stepper motor.
9. The reflector antenna system of claim 1 wherein the ribs are
comprised of composite materials.
10. The reflector antenna system of claim 1 further comprising a
plurality of stowage devices for holding the reflector in the
collapsed and stowed configuration.
11. The reflector antenna system of claim 1 further comprising at
least one clam-shell-type storage clamp for holding the reflector
in the collapsed and stowed configuration.
12. The reflector antenna system of claim 11 wherein said storage
clamp comprises a first set of spherical connectors positioned
between each of said ribs and a second set of spherical connectors
is positioned between each of said ribs and said storage clamp.
13. The reflector antenna system of claim 11 wherein said storage
clamp comprises two sets of spherical connectors positioned between
each of said ribs.
14. The reflector antenna system of claim 11 wherein said storage
clamp is double-acting in order to permit said boom to pass
therethrough during deployment.
15. The reflector antenna system of claim 1 wherein said mesh
member is comprised of a plurality of substantially flat facets,
the corners of said facets being aligned with said ribs.
16. An umbrella-like reflector assembly comprising:
an actuable hub assembly;
a main rib connected to said hub assembly by a hinge mechanism;
a plurality of secondary ribs each connected to said hub assembly
by a respective hinge mechanism;
a mesh member attached to said main rib and said plurality of
secondary ribs for providing a reflective surface; and
a depolyment boom connecting said main rib of the reflector to a
spacecraft, wherein said deployment boom is operable with said main
rib to move the reflector between a collapsed and stowed
configuration adjacent to the spacecraft and a deployed
configuration away from the spacecraft;
wherein activation of said hub assembly causes the reflector to
move between said collapsed and deployed configurations.
17. The reflector assembly of claim 16 wherein the orientations of
said hinge mechanisms are different for each of said ribs such that
the stowed width of the reflector is minimized to permit the
reflectorto fit adjacent said spacecraft within the confines of the
payload fairing.
18. The reflector assembly of claim 16 wherein said main and
secondary ribs are contoured to fit the shape of the reflector.
19. The reflector assembly of claim 16 wherein said secondary ribs
are truss-shaped.
20. The reflector assembly of claim 16 wherein said plurality of
secondary ribs are fabricated from GFRP sandwich plates.
21. A method of forming the surface of a mesh reflector having a
hub assembly and a plurality of ribs, wherein the plurality or ribs
have an inner and an outer portion, the method comprising the steps
of:
optically aligning the outer portions of each of said plurality of
ribs;
optically aligning the hub assembly and inner portions of each of
the plurality of ribs;
splicing the outer portions of each of the plurality of said ribs
to the respective inner portions;
installing a mesh member over said ribs;
installing a network of tensioning chords to said mesh member;
and
attaching said mesh member to said ribs along radial attachment
points on said ribs.
22. The method of claim 21 further comprising the steps of
optically measuring the surface of the mesh reflector.
23. The method of claim 22 further comprising adjusting said mesh
member until the surface of the mesh reflector is satisfactory.
24. The method of claim 21 further comprising the step of
kinematically supporting each of said outer rib portions during
alignment in a total of six degrees of freedom.
25. The method of claim 24 wherein two stand members are utilized
to provide the kinematic support with the locations of the stands
being optimized to minimize rotation of the inner ends of said
outer rib portions where said inner and outer rib portions are to
be spliced together.
26. The method of claim 24 wherein two stand members are utilized
to provide the kinematic support with the locations of the stands
being optimized to minimize deflection of the tooling points on the
ribs used for alignment.
27. The method of claim 21 further comprising the step of
preloading said inner rib portions at the time of splicing them to
said outer rib portions, with forces and/or moments equivalent to
the loads expected to be imparted to them in space by said outer
rib portions when at least said mesh member is attached thereto and
preloaded appropriately.
28. The method of claim 21 further comprising the step of reducing
the contours of said ribs by an amount substantially equivalent to
the predicted deflections expected to be imposed to them in space
due at least to said mesh member.
Description
TECHNICAL FIELD
The present invention relates to deployable satellite reflector
antennas, and more particularly, to an edge-supported collapsible
mesh type reflector antenna of the type launched and sustained in
space.
BACKGROUND ART
High gain antenna reflectors have been deployed into space for
several decades. The configurations of such reflectors have varied
widely as material science has developed and as the sophistication
of technology and scientific needs have increased.
Large diameter antenna reflectors pose particular problems during
all phases of their existence, whether it is assembly, stowage,
launch, deployment and/or usage. Double-curved, rigid surfaces
which are sturdy when in a deployed position cannot be easily
folded for storage. Often, reflectors are stored a year or more in
a folded, stowed position prior to deployment. In an attempt to
meet this imposed combination of parameters, large reflectors
sometimes have been segmented into petals so that these petals
could be stowed in various overlapping configurations. However, the
structure required in deploying such petals has tended to be rather
complex and massive, thus reducing the practical feasibility of
such structures. For this reason, dish-shaped antenna reflecting
surfaces larger than those that can be designed with petals
typically employ some form of a compliant structure.
Responsive to the need for such a compliant structure rib and mesh
designs have been supplied and utilized. A network of tensioned
radial and circumferential chords divides the mesh into
substantially flat facets. The effect on the reflector performance
caused by the difference in shape between these flat facets and the
true parabolic surface is referred to as the faceting error. Prior
art mesh reflector designs require the use of numerous facets
because the circumferential and angular spacing between the ribs
and the mesh attachment locations are not optimized to minimize the
faceting error.
Other antenna designs typically include a center post about which
the petals are configured, much like an umbrella configuration.
This also affects the reflective quality of the resulting surface,
because the center portion typically is the point of optimum
reflectance, which is often blocked by the center post. Thus, it is
desirable to have a structure that is deployable from a compact,
stored position to an open dish-shaped position without center post
blockage.
More recently, many rigid antenna reflectors have been constructed
from graphite fiber reinforced, plastic materials (GFRP). Such
materials may satisfy the requirements for space technology and
contour accuracy and, therefore, high performance antenna systems.
However, power and performance of rigid antennas are limited, owing
to the size of the payload space in a launch vehicle. Very large
completely rigid antennas are highly impractical to launch into
space, hence the requirements for practical purposes can be
satisfied only when the antenna is of a collapsible and foldable
construction.
At present, antenna reflectors of the collapsible and foldable
variety are of two design types. One type is a grid or mesh-type
reflector that is folded like an umbrella. The other type includes
foldable rigid and hinged petal members. Antennas of the second
type are available in a variety of configurations, some of which
are disadvantaged by the requirement for an excessive number of
joints and segment pieces which, owing to the particular folding
and collapsing construction, are of different shape and size. Also,
the larger the number of hinges and segments, the more complex will
be the deployment mechanism and its operation. Any added weight
also is a disadvantage relative to a satellite system.
For a given paraboloid reflector diameter, the number of ribs used
determines the width of each mesh singly-curved gore. Thus, more
ribs result in more and narrower mesh gores, with each narrower
gore being a better approximation of the ideal paraboloid shaped
gore.
While the existing paraboloid reflectors are satisfactory to some
degree, they have several inherent disadvantages which detract from
their usefulness. Among the foremost of these disadvantages are
excessive weight, excessive stowage volume requirements, excessive
cost and complexity, inadequate surface accuracy, and inadequate
deployment reliability.
SUMMARY OF THE INVENTION
Accordingly, it is an object of the present invention to provide an
improved umbrella-type reflector having a low stowage profile. It
is also an object of the present invention to provide a mesh-type,
dish-shaped reflector which is an improvement over known mesh-type
reflectors.
It is another object of the present invention to provide an
edge-supported mesh-type umbrella reflector having a main rib
supported by a boom on a spacecraft. It is still another object of
the present invention to provide an edge-supported umbrella
reflector which may be fed by an offset feed assembly.
It is still yet another object of the present invention to provide
a mesh-type edge-supported umbrella reflector having a main rib and
a plurality of secondary ribs each connected to a hub assembly by a
single hinge without additional mid-rib hinges. It is a further
object of the present invention to provide an edge-supported
umbrella reflector having a mesh member attached to the ribs and
uneven circumferential spacing between the ribs to minimize the
faceting error of the reflector.
It is still a further object of the present invention to provide an
edge-supported umbrella reflector having uneven radial spacing
between the attachment points of the mesh to minimize the faceting
error of the reflector. It is still yet a further object of the
present invention to provide a mesh-type edge-supported umbrella
reflector having hinge axis orientations for the ribs optimized to
effect the tightest possible folding to achieve a low stowage
profile.
In carrying out the above objects and other objects, features, and
advantages of the present invention, a mesh-type umbrella-like
reflector for use on an orbiting spacecraft is provided. The
reflector has a contoured main rib and a plurality of contoured
secondary ribs each connected to a hub assembly by a respective
hinge such that activation of the hub assembly causes the reflector
to move between collapsed and opened configurations. The mesh
member is attached to the ribs. A deployment boom connects the main
rib of the reflector to the spacecraft. The deployment boom is
operable with the main rib and the spacecraft to move the reflector
between a stowed configuration proximate the spacecraft and a
deployed configuration outside the spacecraft.
A feed assembly is connected to the spacecraft. The feed assembly
is offset from and operable with the mesh member of the reflector
when the reflector is in the opened and deployed configurations to
receive and/or transmit radio frequency energy therefrom.
The advantages of the present invention are numerous. For example,
the reflector stowed profile is sufficiently slim to permit the
stowage of a reflector up to twenty-five meters in diameter
attached to a full-sized spacecraft (via two or more clam-shell
type deployable clamps) on one or more commercially available
launch vehicles. These and other features, aspects, and embodiments
of the present invention will become better understood with regard
to the following description, appended claims, and accompanying
drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 illustrates the overall arrangement of a preferred
embodiment of the mesh-type edge-supported umbrella reflector of
the present invention;
FIG. 1A illustrates the edge-supported umbrella reflector in a
stowed configuration within a booster payload fairing of a
spacecraft;
FIG. 2 illustrates a hinge connecting a rib to a hub assembly;
FIGS. 3a-3d illustrate the deployment sequence of the
edge-supported umbrella reflector;
FIGS. 4a-4c illustrates an exemplary launch constraint clamp;
FIGS. 5a-5c illustrate the deployment sequence of the hub
assembly;
FIG. 6 illustrates the hub assembly restraining a secondary rib
against deployment;
FIG. 7 illustrates the mesh layout of the edge-supported umbrella
reflector;
FIGS. 8-13 illustrate the construction of the main rib; and
FIGS. 14-15 illustrate the construction of one of the secondary
ribs.
BEST MODES FOR CARRYING OUT THE INVENTION
Referring now to FIG. 1, an edge-supported umbrella reflector
assembly 10 of the present invention is shown. Umbrella reflector
assembly 10 includes a reflector 12 connected to a spacecraft 14 by
a relatively stiff deployment boom 16. Reflector 12 is shown in
FIG. 1 in a deployed configuration and shown in dashed lines in a
stowage configuration within a booster payload fairing 15.
Reflector 12 may have a diameter ranging from six to twenty-five
meters. FIG. 1 shows a 15.times.12.3 meter reflector in a mid-sized
booster fairing.
FIG. 1A illustrates reflector 12 in a stowed configuration within
booster payload fairing 15. Booster payload fairing 15 shown in
FIG. 1A is a Long March III-B fairing.
Reflector 12 includes a main rib 18 and a plurality of secondary
ribs 20. Boom 16 connects main rib 18 to spacecraft 14. Main rib 18
has a torque box construction and contoured edges and is connected
to boom 16 to provide an "edge-support" for reflector 12. Secondary
ribs 20, which are described in more detail below, are of a light
weight planer-truss construction and are contoured and tapered
toward their outer edges.
Attached to main rib 18 and secondary ribs 20 is a mesh 22 which
acts as a reflecting surface. Reflector 12 further includes a hub
assembly 24. Hub assembly 24 is connected to main rib 18 and
secondary ribs 20 and assists in moving the ribs between the
deployed and stowed configurations. A feed assembly 26 on
spacecraft 14 is operable with reflector 12 to transmit and/or
receive radio frequency (RF) energy therefrom. Feed assembly 26 is
offset from the edge of reflector 12 thus avoiding self blockage by
the feed assembly of the reflected antenna RF energy.
A first deployment actuator 28 connects a top end 30 of boom 16 to
main rib 18. A second deployment actuator 32 connects a bottom end
34 of boom 16 to spacecraft 14. Deployment actuators 28 and 32 are
preferably of the conventional viscous damped spring actuator
type.
To minimize the width of reflector 12 at the critical location in
the spacecraft payload compartment, shown by reference numeral 39
in FIG. 1A, where the stowed reflector passes between one edge 42
or spacecraft shelf 36 and booster payload fairing 15, the
deployment boom is kinked at that location as shown by reference
numeral 38.
A pair of secondary ribs 20a and 20b generally opposite of main rib
18 are spaced apart when stowed to permit the passage and nesting
of boom 16 between them on the same plane and opposite to the main
rib. In FIG. 1, rib 20a falls directly behind rib 20b due to
symmetry and is not specifically shown in the Figure. The number of
secondary ribs 20 is an even number such that the total number of
ribs 18 and 20 (and thus the number of triangular reflector gore
segments) is an odd number. Thus, none of secondary ribs 20 falls
directly opposite main rib 18, where boom 16 stows.
FIG. 2 illustrates the connection of a rib such as secondary rib 20
to hub assembly 24. Each of ribs 18 and 20 is connected by a single
hinge to hub assembly 24. For instance, as shown in FIG. 2, hinge
40 attaches secondary rib 20 to hub assembly 24. The hinges are
designed to be zero-clearance (pre-loaded) hinges. The hinge
construction shown is aimed at minimizing the center spacing
between the hinges, and thus the diameter of hub assembly 24, while
permitting rib assembly and disassembly. The small hub diameter
(about 4% of the reflector diameter) permits stowage of reflector
12 in the often unused volume near the top 41 of booster payload
fairing 15.
The hinge axis orientations for each of the ribs are individually
optimized to effect the tightest possible folding thus minimizing
the width of reflector 12 where the reflector passes between
spacecraft corner 42 and the booster payload fairing 15 without
significantly compromising the width of the reflector in the
orthogonal direction. The orthogonal direction is the direction
perpendicular to the view shown in FIG. 1.
Referring now to FIGS. 3a-3d, the deployment sequence which
reflector 12 performs on orbit to transition from the stowed launch
configuration to the operational deployed configuration is shown.
FIG. 3a illustrates reflector 12 in the stowed and launch
configurations. A plurality of stowage clamps 46(a-c) hold
reflector 12 to spacecraft 14. Stowage clamps 46(a-c) include
pyrotechnic devices (e.g., bolt cutters or separation nuts) to lock
and release the stowage clamps as will be explained later with
reference to FIG. 4.
During the first motion of deployment shown in FIG. 3b, pyrotechnic
devices on stowage clamps 46(a-c) are released permitting the
activation of first deployment actuator 28 connecting main rib 18
to boom 16. First deployment actuator 28 causes reflector 12 to
move away from spacecraft 14 as shown in FIG. 3b.
During the second motion of deployment shown in FIG. 3c, a launch
lock 48 attaching a point near the kink location 38 of boom 16 to
lower stowage clamp 46c is released. Correspondingly, second
deployment actuator 32 connecting bottom end 34 of boom 16 to
spacecraft 14 is activated. Second deployment actuator 32 causes
reflector 12 to move up and around spacecraft 14 as shown in FIG.
3c. As can be seen, this motion passes boom 16 through the upper
stowage clamp 46a which is facilitated by the particular design of
the clamp to be discussed in relationship to FIG. 4.
FIG. 3d depicts reflector 12 in the operational deployed
configuration. To achieve the deployed configuration from the
second motion of deployment shown in FIG. 3c, hub assembly 24 is
activated to force ribs 18 and 20 open relative to hub assembly 24
as will be explained in greater detail with reference to FIG. 5.
Accordingly, in the deployed configuration, reflector 12 is
operational with offset feed assembly 26 to transmit and/or receive
RF energy therefrom. If desired, a second reflector assembly 50 on
spacecraft 14 may be employed for a different frequency band in
addition to reflector 12.
FIG. 4 illustrates an exemplary stowage device 46. The stowage
device 46 is double acting with both front half 52 and back half 54
deployable in order to permit passage of boom 16 through the
stowage device during the second motion of deployment described
above. Front half 52 and back half 54 includes respective arms
56(a-b) and 58(a-b). Arms 56(a-b) and 58(a-b) are pivotable about a
respective hinge assembly 51(a-d) with an associated
crushable/catcher fixture assembly 60 and 62. Arms 56a and 58a are
connected by a separation bolt having a bolt cutter 64 and a bolt
catcher 66. The separation bolt is releasably engaged to allow arms
56a and 58a to open. Release is accomplished via bolt cutter 64
which is pyrotechnically operated using small explosive charges to
sever the separation bolt upon ground command. Other pyrotechnic
devices such as separation nuts may be used alternatively to
perform this function. Arms 56b and 58b are similarly arranged.
Arms 56 and 58 include adjustable screws 53 having hemisperical
heads which engage dry lubricated metallic washers with sperical
indentations 55 bonded (or otherwise attached) to each of secondary
ribs 20 and main rib 18. Additionally, at the stowage device
locations, ribs 18 and 20 are spherically rotatably engaged to each
other using pairs of male spherical protrusions 57 and dry
lubricated female washers with sperical indentations 59 attached to
the ribs via light weight stand offs (61a,b). In certain locations,
it may not be practical (or desirable) to have a direct connection
between the stowage device and ribs 20, as is the case with ribs
20c and 20d where such an attachment may impede deployment (because
the stowage device at these locations does not move). For such
ribs, additional sets of spherical attachments 57, 59 connecting
ribs 20(c-d) to their neighborhing ribs at locations marked 63 are
used instead of connections to the stowage device.
Referring now to FIGS. 5a-5c, deployment of hub assembly 24 to move
reflector 12 into the deployed configuration is shown. Hub assembly
24 includes a hub 67. A shaft 68 and two stepper motors 70(a-b) are
connected to hub member 67. Hub assembly 24 further includes a base
plate 72. A motor strap 72 wraps around pullies 76(a-b) connected
to base plate 72 and connects at its two ends to pullies mounted to
respective stepper motors 70(a-b). In the stowed configuration
shown in FIG. 5a, base plate 72 restrains secondary rib 20 against
deployment by engaging the secondary rib through a shear code 77
shown in greater detail in FIG. 6.
Secondary rib 20 is connected to hub member 67 by hinge 40. A lower
heavy GFRP strap 78 connects secondary rib 20 to base plate 72. A
relatively flexible upper strap 80 connects secondary rib 20 to
shaft 68 above hub 67. Deployment of reflector 12 is effected by
activating either or both of stepper motors 70(a-b) operatively
with motor strap 74, pullies 76(a-b), and base plate 72 to
redundantly slowly drive shaft 68 upwards through hub 67. As shaft
68 travels upwards, upper strap 80 pulls on secondary rib 20
causing it to extend away from hub assembly 24 as shown in FIG. 5b.
When shaft 68 is in a fully deployed position extending above hub
67, base plate 72 is completely behind the theoretical reflector
surface and reflector 12 is in the deployed configuration shown in
FIGS. 1 and 5c.
Deployment of reflector 12 and hub assembly 24 is terminated by the
engagement of at least one of two redundant spring-loaded detents
into holes located in shaft 68 such that they line up with the
detents when reflector 12 is in the deployed configuration (not
specifically shown). It should be noted that while in FIGS. 5a-5c
and in the discussion above, hub member 67 is represented as
stationary with ribs 20 and shaft 68 moving relative to it. In
reality, hub member 67 rotates approximately 90 degrees during the
phase of deployment as can be seen from comparing FIGS. 3c and 3d.
This slow rotation is a rigid body motion and does not affect the
kinematics of deployment nor the relative motions between the
various components described above.
Hub assembly 24 is capable of slowly controlled (non-dynamic),
reversible deployment in 1-G environment without off loading
(except for main rib 18) initiated without irreversible pyrotechnic
events. Hub assembly 24 incorporates all moving parts into a
compact separately testable assembly, thus maximizing deployment
reliability and testability.
FIG. 7 illustrates the layout of mesh member 22 on reflector 12.
Mesh member 22 is divided into a plurality of trapezoidal-shaped
facets 82 by a network of pre-tensioned Kevlar or Vectran radial
chords 84 and circumferential chords 68. Chords 84 and 86 are
constructed on the focus side (towards feed assembly 26) of mesh
member 22. Mesh 22 is thus divided into substantially flat facets
82. Mesh 22 is attached to ribs 18 and 20 only at corners 88 of
facets 82. In short, mesh 22 is attached at radial attach points
running along ribs 18 and 20. The effect on the performance of
reflector 12 caused by the difference in shape between flat faces
82 and the true parabolic surface is referred to as the faceting
error.
For a given diameter of reflector 12, the number of reflector ribs
is chosen to limit the faceting error to an acceptable value. In
the present invention, the faceting error resulting from a given
number of ribs, or conversely, the number of ribs required to limit
the faceting error to a given level, is further optimized by three
characteristics.
First, the circumferential spacing between adjacent ribs 18 and 20
is varied across reflector 12. For reflector 12 fed by offset feed
assembly 26, the vertex of the reflector is near the outer end of
main rib 18 where it connects to first deployment actuator 28. The
curvature of reflector 12 is the highest nearest the vertex.
Accordingly, main rib 18 and adjacent secondary ribs 20 have a
higher curvature than secondary ribs 20 farthest away from the
vertex. Pair of secondary ribs 20(a-b) opposite from main rib 18
have the lowest curvature. The circumferential spacing between the
rib tips is reduced for the ribs nearest the vertex and gradually
increases as the ribs extend to the opposite end near ribs 20(a,b).
Thus, secondary ribs 20(a-b) have the largest angular spacing and
secondary ribs 20 adjacent on each side of main rib 18 are spaced
from the main rib with the smallest circumferential spacing. The
purpose of using uneven spacing between ribs 18 and 20 is to
approximately equalize the normal distance between the outermost
circumferential chords and the parabolic surface.
Second, the number of radial attachment points of mesh member 22
along ribs 18 and 20 are appropriately selected. For instance, it
can be shown that if the objective is to minimize the total number
of radial attachment points then the optimum number of radial
attachment points is equal to the number of ribs divided by (.pi.
multiplied by the square root of 2): ##EQU1## However, because the
number of radial attachment points has significantly less impact on
cost and weight of reflector 12 than the number of ribs, the number
of radial attachment points is selected to be at least equal to the
number of ribs divided by .pi..
Third, the radial spacing between the radial attach points along
ribs 18 and 20 decreases as the circumference of reflector 12
increases. Because the faceting error is proportional to the area
of the facet multiplied by the square of the maximum distance from
the facet to the parabolic surface and by the power density of the
feed illumination (B), optimum spacing between the radial attach
points is achieved when the quantity (W*L*(W.sup.2 +L.sup.2).sup.2
*B) is approximately equal for all facets. W and L are the average
width and length of a facet, respectively. The phase relationships
between the various radiating feed elements of feed assembly 26 are
also optimized to minimize the faceting errors.
Referring now to FIGS. 8-13, the construction of main rib 18 is
shown. Main rib 18 consist of two portions. Namely, inner main rib
90, which starts out as part of hub assembly 24, and outer main rib
92. Ribs 90 and 92 each have a bonded built-up box beam
cross-section and is fabricated primarily from GFRP plates, angle
members, and channel members. Outer main rib 92, including its
integral end fitting 94 is fabricated primarily from only two
different thickness plates 95 and 96, one channel member 97, and
four different size angle members 98(a-d). The curved reflector
contour of outer main rib 90 is provided by numerically controlled
(N/C) machining of the side plates to the required profile. Tooling
holes 99 are provided in the side plates and near the ends of each
channel and angle to aid in assembling main rib 18.
Referring now to FIGS. 14 and 15, the construction of one of
secondary ribs 20 is shown. Secondary rib 20 consists of an inner
secondary rib 110, which is a part of hub assembly 24, and an outer
secondary rib 112. Because there is a relatively large number of
secondary ribs, they account for the largest single weight item of
reflector 12. It is therefore important to design the secondary
ribs with a low cost, light-weight structure. Specifically,
secondary rib 20 has a planar truss (frame) shape N/C machined (or
waterjet cut) from a large honeycomb sandwich plate. The sandwich
plate has thin GFRP facesheets and a non-metallic core made of
Nomex, Corex or Kevlar. Depending on the size of reflector 12 and
the machining facility available, outer secondary rib 112 is made
out of one to three segments spliced together using small bonded
GFRP doubler plates with the aid of simple flat tooling with
indexing tooling holes/pins. This approach minimizes fabrication
time and tooling cost, permits maximum flexibility for rib weight
optimization, and provides an accurate contour shape. The absence
of mechanical joints (except for the one preloaded hinge per rib)
and the minimum number of bonded joints makes for a highly
predictable structural and thermo-structural behavior for reflector
12.
Due to the extremely favorable specific strength and stiffness
characteristics, as well as their low coefficient of thermal
expansion (CTE), composite materials make up over 98% of the volume
of reflector 12. Stepper motors 70(a-b) account for more than half
the weight of the tiny amount of metallic materials used, with the
remainder of the weight confined to small components such as
fasteners, monoballs, bushings, etc., which have no detrimental
effect on thermal distortion.
For weight and material cost efficiency, the choice of the type of
graphite fiber used for secondary ribs 20 is important. Because the
design is generally stiffness and/or stability driven, the cost per
unit stiffness is the most significant parameter in order to
minimize cost. The specific compressive stiffness is the preferred
measure for stiffness and stability efficiency. Ultra-high modulus
Graphite fibres of Toray Industries, Inc. designated as M55J, have
a low cost per unit stiffness. Ultra-high modulus Graphite fibres
of Nippon Graphite Fibre corporation designated as XN70 have a high
specific compressive stiffness. Accordingly, M55J is preferably
used for the construction of secondary ribs 20 because it has a
specific compressive stiffness of 85% of that of XN70 at less than
half the cost per pound, as well as significantly higher
strength.
The present invention provides a reflector assembly having
maximized deployment reliability and performance with minimum cost.
High surface accuracy resulting in high reflector performance is
enhanced by two general features. First, enhanced deployment
repeatability and second, minimum thermal distortion.
Deployment repeatability is enhanced by two specific features.
First, pre-loaded monoballs or ball bearings 41 are used to form
the rib/hub hinges 40. Hinges 40 are further pre-loaded by use of
two sets of deployment straps 78 and 80 which results in a
repeatable hinge contact point regardless of the magnitude of the
tension in either strap. This enhances repeatability by eliminating
the effect of hinge sloppiness on the deployed shape.
Second, mechanical contact-type deployment stops are eliminated.
Instead, heavy GFRP straps 78 permanently connecting ribs 18 and 20
to base plate 72 are used as the stops. Straps 78 have very high
axial stiffness and very low CTEs. In contrast, conventional
mechanical stops are often metallic (high CTE), may exhibit local
yielding (thus have low apparent stiffness), and may contact at
slightly different points at successive deployments resulting in
shape changes (non-repeatability in deployment).
Thermal distortion is minimized by these specific techniques.
First, the choice of the composite lay-ups is selected consistent
with the type of graphite fiber used. The result is very low CTEs
in the range of +0.05 to -0.20 ppm/degree F. The addition of minor
amounts of adhesive, foam fill, and/or metallic fasteners/inserts,
result in effective CTEs in the range of +0.1 to -0.2 ppm/degree
F., which in turn minimizes thermal distortion.
Second, thermal blankets are positioned around ribs 18 and 20 and
boom 16. The thermal blankets reduce the gradient through the
thickness and across the depth of the ribs and the boom further
reducing the thermal distortion. The blankets, which are designed
to be fabricated using pressure sensitive adhesive rather than
Velcro tape, serve the additional function of protecting mesh
member 22 from possible snagging on exposed honeycomb core edges or
fastener heads.
Third, the effect of the relatively high mesh CTEs is rendered
negligible by the use of an extremely low stiffness tricot knit.
The effect of the moderately low CTE and moisture sensitivity of
the Kevlar mesh retension chords is also rendered negligible by the
use of soft springs in series with each of these chords.
High deployment reliability and low cost is also achieved by other
general features. First, reflector 12 is deployable in one-G
without the need for off-loading ground support equipment (GSE).
Only main rib 18 is off-loaded using dead weights and pulleys
hanging from a crane or by using a helium-filled balloon. This
avoids the expense, facility limitation, and errors and
uncertainties induced by a huge multitrack off-loader system. The
high capability deployment system required to accomplish the task
in 1-G will provide a high deployment margin in zero-G. In addition
to the increased deployment system capability, the 1-G
deployability is made possible by the highly efficient rib
structural design (high stiffness, super lightweight trussed
graphite honeycomb) and the ultra lightweight mesh and mesh
restraint chords utilized.
Second, slow (non-dynamic), motorized, and reversible deployment by
hub assembly 24 without pyrotechnic initiation results in
significant reliability and cost advantages. These provide the
ability to overcome deployment hang ups by backing up the motors
for a certain distance and then re-deploying. This minimizes or
eliminates the need for expensive dynamic deployment analysis and
its associated uncertainties. The invention is also deployable in
ambient 1-G environment (no air drag effect or gravity/off loader
induced drag error during ground testing), without any pyroshock,
pyro refurbishment, or associated adverse reliability impact.
Third, soft tooling integration concept produces high surface
accuracy at low cost. The soft tooling integration concept
eliminates the cost and facility requirements associated with large
tooling required for typical spacecraft reflectors of the prior
art. The soft tooling integration concept involves several steps
labeled A-G in the following paragraphs.
A) Each outer rib is supported in a kinematic fashion (at a total
of six degrees of freedom) at two locations and optically aligned
in its theoretical position. The location of the two support points
for each rib is selected to minimize the rib tooling point
deflections and the rotation of the inner end of the outer rib
where it will be subsequently spliced to the inner rid due to 1-G
loading.
B) In defining the outer rib contours, the profiles are cut back
relative to the theoretical parabolic shape by the deflections
predicted to be caused by the nominal mesh and mesh retaining chord
pre-loads for a case where the ribs are fixed at the interface
points between the inner and outer ribs. The error associated with
the uncertainty in this process is minimized by designing the ribs
to be particularly stiff in their own planes (which is also needed
in order to handle 1-G deployment and launch loads).
C) The hub/inner rib assembly is optically aligned to its
theoretical position while supported on a stiff adjustable
stand.
D) The outer edges of the inner ribs are loaded with a set of
pre-calibrated tension springs and moment arms. These loads
represent the forces and moments predicted to be caused by the
nominal mesh and mesh retaining chord pre-loads.
E) After alignment, the outer ribs are spliced to the inner ribs
via field splice joints injected with adhesive and syntactic foam
which act as liquid shims to fill any gaps in the joints between
the edges of the inner and outer ribs without appreciably stressing
them.
F) Mesh and mesh retension chords are installed and tensioned to
their desired levels.
G) Reflector contour is measured (including 1-G
compensation/off-loading) and final contour adjustments are made if
necessary. Adjustments are made either by slight changes in mesh
retension chord tensions, and/or hub strap tensions. Contour
measurements are then repeated and so on until contour shape is
satisfactory.
A 40'.times.50' (12.3 meter projected aperture) engineering
development model reflector utilizing the teachings of the present
invention was designed, built and tested. Photogrammetric surface
measurements were taken showing that the reflector met the as-built
RMS goal of 1 mm. Three successful deployment demonstrations were
performed (two prior to vibration testing and one post-vibration
testing). Moreover, a protoflight level sine vibration test was
performed with the reflector supported on a spacecraft simulation
fixture and successfully completed as indicated by a post-test
functional deployment and surface measurement demonstration.
It should be noted that the present invention may be used in a wide
variety of different constructions encompassing many alternatives,
modifications, and variations which are apparent to those with
ordinary skill in the art. Accordingly, the present invention is
intended to embrace all such alternatives, modifications, and
variations as fall within the spirit and scope of the appended
claims.
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